Combustion gas discharge nozzle for a rocket engine provided with a sealing device between a stationary part and a moving part of the nozzle

Patent metadata

Field	Value
Publication number	US-10316796-B2
Application number	US-201414905923-A
Country	US
Kind code	B2
Filing date	Jul 23, 2014
Priority date	Jul 26, 2013
Publication date	Jun 11, 2019
Grant date	Jun 11, 2019

Abstract

The invention relates to a combustion gas discharge nozzle for a rocket engine including a stationary part and a moving part extending from the stationary part, the moving part made using flaps positioned downstream from the stationary part and forming an extension of the nozzle, the nozzle including a sealing device extending between the fixed part and the moving part in the form of a flexible membrane withstanding a local temperature of the combustion gases at the nozzle outlet and connecting the end of the stationary part to a border of the flaps or petals forming the moving part, the flexible membrane forming an annular tubing, the sealing device being provided with a duct for injecting gas at the flexible membrane between the stationary part and the moving part extending the nozzle.

First claim

What is claimed is:

1. A nozzle for ejecting combustion gas from a rocket engine comprising:
a fixed part;
a moving part extending the fixed part, said moving part produced in the form of flaps positioned downstream of the fixed part and forming an extension of the nozzle, wherein the nozzle further comprises:
a flexible membrane providing sealing between the fixed part and the moving part, the flexible membrane being resistant to a local temperature of the combustion gas leaving the nozzle, the flexible membrane connecting a downstream end of the fixed part to a border of the flaps that form the moving part, the flexible membrane forms an annular pipe provided with a duct for conveying an exhaust gas from a turbine of a turbopump to the flexible membrane, the exhaust gas being injected between the fixed part and the moving part extending said nozzle.

2. The nozzle as claimed in claim 1 , wherein the annular pipe is configured to distribute the exhaust gas over a perimeter of an outlet section of the fixed part of the nozzle.

3. The nozzle as claimed in claim 1 , wherein the annular pipe is situated in line with a rocker arm of the moving part.

4. The nozzle as claimed in claim 1 , wherein the flexible membrane is silica-based fabric able to withstand continuously a temperature of at least 1000° C. (1832° F.).

5. The nozzle as claimed in claim 1 , wherein the flexible membrane comprises a ceramic insulator between two fabrics, one of one of the two fabrics comprising a fabric made from an aluminoborosilicate refractory ceramic fiber on a hot face side of the flexible membrane, the other one of the two fabrics comprising an aramid fiber type such as poly-paraphenylene terephthalamide on a cold face side of the flexible membrane, the aramid fiber type fabric being intended to give the flexible membrane mechanical integrity.

6. A rocket engine comprising the nozzle with the fixed part, the moving part and flexible membrane as claimed in claim 1 , wherein the flaps are arranged around an exit section of the rocket engine nozzle as an extension of the fixed part.

7. The rocket engine as claimed in claim 6 , wherein the flexible membrane forming the annular pipe for injecting the exhaust gas from the turbine of the turbopump of the rocket engine between the fixed part of the nozzle and the moving part extending said nozzle, a pressure of the exhaust gas being regulated so as to be higher than a pressure of the combustion gas leaving the fixed part of the nozzle.

8. An aircraft comprising a rocket engine as claimed in claim 6 , wherein the flaps are articulated on a frame of the aircraft.

9. The aircraft as claimed in claim 8 , wherein the frame forms part of a rear fuselage in which the rocket engine is installed.

10. The aircraft as claimed in claim 8 , wherein the flaps are able to move and able to allow the moving part to adopt at least one of:
a closed cone shape, so as to provide an aerodynamic external shape that minimizes drag of an aft end of the aircraft during atmospheric flight with the rocket engine not lit;
a cylinder shape at the moment of ignition of the rocket engine; and
a conical divergent shape in the extension and continuation of the exit section of the rocket engine nozzle so as to encourage expansion of the rocket engine combustion gas.

11. The aircraft as claimed in claim 10 , wherein the openness of the conical divergent shape can be varied according to an increase in altitude during rocket engine propelled flight.

12. The aircraft as claimed in claim 10 , further comprising at least one link rod for pivoting the flaps which is able to allow differential openings or closings of said flaps which deflect a jet of the rocket engine combustion gas and thus create a lateral thrust component that allows the aircraft to be steered about a pitch axis and a yaw axis.

13. The aircraft as claimed in claim 8 , wherein the flaps are arranged in two rows with internal/external surfaces of adjacent flaps overlapping one another and are suited to allowing a variation in the exit section of the nozzle while at the same time maintaining an overlap that minimizes leaks of combustion gas between the two rows.

Assignees

• Airbus Defence & Space Sas
• Airbus Safran Launchers Sas

Inventors

• Dutheil Jean-Philippe

Classifications

• F05D2240/128

  Nozzles · CPC title

• F02K1/805 Primary

  Sealing devices therefor, e.g. for movable parts of jet pipes or nozzle flaps · CPC title

• F02K9/978

  Closures for nozzles; Nozzles comprising ejectable or discardable elements · CPC title

• F02K9/97

  Rocket nozzles (thrust or thrust vector control F02K9/80) · CPC title

• F05D2220/80

  in supersonic vehicles excluding hypersonic vehicles or ram, scram or rocket propulsion · CPC title